Novel leading edge film cooling experiments and geometries - Holgate, NE (2019)
The continuing rapid growth of civil aviation and the associated greenhouse gas emissions necessitate further reduction of gas turbine engine specific fuel consumption. Since the use of high pressure compressor bleed air to cool combustor and turbine components is parasitic to engine efficiency, one of the key drivers of this progress is the ongoing improvement of the efficiency with which the coolant air is used. Because it is exposed to the undiluted and highly turbulent combustor outlet gases, the leading edge of the nozzle guide vane requires a sizeable proportion of the overall turbine coolant. Moreover, the low mainstream flow momentum in the leading edge stagnation region means that film coolant jets are highly prone to lift off of the surface, so are often poorly utilised. This study investigates the performance of this film coolant both experimentally and computationally, and proposes a novel film cooling geometry in which jets initially travel against the mainstream flow before crossing the geometric stagnation line. This is shown to achieve vastly more effective film coolant usage near the stagnation line than can conventional designs, without compromising performance farther downstream on the vane surface. Additionally, the importance of combustor dilution port flows in determining the performance of conventional leading edge film cooling designs is quantified, as is the importance of infrared calibrations which take into account temperature inhomogeneities which can occur in advanced aerothermal experimental apparatuses. A novel infrared thermography technique has been developed for determining both of the ubiquitous film cooling performance parameters in a single experiment, where usually two are required. Also proposed is a leading edge coolant supply passage geometric modification which reduces the risk of destructive ingestion of hot combustion gases into film cooling holes. This innovation is also conducive to the improvement of gas turbine engine efficiency, primarily through its potential to facilitate reduced parasitic combustor pressure losses.